It has long been known in gas turbines to provide a dilution air zone in the combustor immediately downstream of the flame zone. The dilution air zone is conventionally located directly within the combustion annulus downstream of the fuel injectors but well upstream of the outlet of the combustor. Generally speaking, dilution air is injected into the combustion annulus defining the flame zone to control the temperature of hot gases.
More specifically, upstream of the dilution zone both fuel and air are injected and ignited in the combustion annulus. It is also conventional for there to be a cooling air film introduced along the walls of the combustion annulus upstream of the dilution zone. Of course, the hot gases that result from combustion of fuel and air then pass toward turbine blades.
As is known, it is important to be able to control the temperature of the hot gases as they enter the nozzle on their way to the turbine blades. This has conventionally been handled by injecting dilution air into the hot gases well upstream of the outlet of the combustor in order to ensure thorough mixing and cooling prior to entry into the nozzle. While effective, this means of controlling the temperature of the hot gases is not satisfactory in every respect.
More particularly, the need to provide the dilution zone in the combustion annulus upstream of the outlet of the combustor tends to dictate the geometry. In other words, the length of the turbine is controlled to a significant degree by the necessity of having a distinct dilution zone within the combustion annulus, i.e., there has been no available manner for satisfactorily shortening the length of the combustor in order to reduce weight and expense. However, conventional designs have also failed to address still another serious problem recognized by those in this field.
In particular, the dilution air flow path is known to cool only a portion of the walls of the combustor. Thus, in a conventional annular combustor of a gas turbine, not only is it true that not all portions of the walls of the combustor are cooled by the dilution air, but the point of injection into the dilution zone has rendered it impossible to effect any significant cooling of the turbine shroud and, thus, of the nozzle and turbine blades. As a result, it has remained to provide a low cost, simple, reliable means of turbine shroud cooling.
As will be appreciated, these problems lead to adverse consequences relative to performance and life span. In other words, due to the heretofore recognized inability to provide an ultra-short combustor and a well-cooled turbine shroud, it has been impossible to achieve the levels of power and fuel economy as well as longer life for the various components such as the nozzle blades, turbine shroud, turbine blades, turbine exhaust duct, etc. Furthermore, if an ultra-short combustor could be provided, there would be less exhaust noise thereby reducing silencing problems.
The present invention is directed to overcoming the above-stated problems by providing a unique gas turbine annular combustor with radial swirling, circumferentially uniform, dilution air injection in a radial flow turbine While the invention has been described in connection with a radial flow turbine, it should be appreciated that the invention could be utilized with any gas turbine construction.